This invention relates generally to gas turbine engines, and in particular, to a process for cooling a flow path surface region on a turbine airfoil.
In gas turbine engines, for example, aircraft engines, air is drawn into the front of the engine, compressed by a shaft-mounted rotary-type compressor, and mixed with fuel. The mixture is burned, and the hot exhaust gases are passed through a turbine mounted on a shaft. The flow of gas turns the turbine, which turns the shaft and drives the compressor and fan. The hot exhaust gases flow from the back of the engine, driving it and the aircraft forward.
During operation of gas turbine engines, the temperatures of combustion gases may exceed 3,000xc2x0 F., considerably higher than the melting temperatures of the metal parts of the engine aft of the compressor, which are in contact with these hot gases. Operation of these engines at gas temperatures that are above the metal part melting temperatures is a well established art, and depends in part on supplying a cooling fluid to the outer surfaces of the metal parts through various methods. Metal parts of these engines that are particularly subject to high temperatures, and thus require particular attention with respect to cooling, are, for example, combustor liners and the metal parts located aft of the combustor including high pressure turbine airfoils, such as turbine blades and turbine vanes.
The hotter the turbine inlet gases, the more efficient is the operation of the jet engine. There is thus an incentive to raise the turbine inlet gas temperature. However, the maximum temperature of the turbine inlet gases is normally limited by the materials used to fabricate the components downstream of the combustors such as the vanes and the blades of the turbine. In current engines, the turbine vanes and blades are made of nickel-based superalloys, and can operate at temperatures of up to 2100xc2x0-2200xc2x0 F. with appropriate well-known cooling techniques.
The metal temperatures can be maintained below their melting levels with current cooling techniques by using a combination of improved active cooling designs and thermal barrier coatings (TBCs). For example, with regard to the metal blades and vanes employed in aircraft engines, some cooling is achieved through convection by providing passages for flow of cooling air from the compressor internally within the blades so that heat may be removed from the metal structure of the blade by the cooling air. Such blades have intricate serpentine passageways within the structural metal forming the cooling circuits of the blade.
Small internal orifices have also been devised to direct this circulating cooling air directly against certain inner surfaces of the airfoil to obtain cooling of the inner surface by impingement of the cooling air against the surface, a process known as impingement cooling. In addition, an array of small holes extending from a hollow core through the blade shell can provide for bleeding cooling air through the blade shell to the outer surface where a film of such air can protect the blade from direct contact with the hot gases passing through the engines, a process known as film cooling.
In another approach, a TBC is applied to the turbine blade component, which forms an interface between the metallic component and the hot gases of combustion. The TBC includes a ceramic coating that is applied to the external surface of metal parts to impede the transfer of heat from hot combustion gases to the metal parts, thus insulating the component from the hot combustion gas. This permits the combustion gas to be hotter than would otherwise be possible with the particular material and fabrication process of the component.
TBCs include well-known ceramic materials, such as, for example, yttrium-stabilized zirconia (YSZ). Ceramic TBCs usually do not adhere well directly to the superalloys used as substrate materials. Therefore, an additional metallic layer called a bond coat is placed between the substrate and the TBC. The bond coat may be made of an overlay alloy, such as a MCrAlX, or other composition more resistant to environmental damage than the substrate, or alternatively, the bond coat may be a diffusion nickel aluminide or platinum aluminide. The surface of the bond coat oxidizes to form a thin, protective aluminum oxide scale that provides improved adherence to the ceramic top coatings. The bond coat and overlying TBC are frequently referred to as a thermal barrier coating system.
Multi-layer coatings are well known in the art. For example, U.S. Pat. No. 5,846,605 to Rickerby et al. is directed to a coating having a plurality of alternate layers having different structures that produce a plurality of interfaces. The interfaces provide paths of increased resistance to heat transfer to reduce thermal conductivity.
Rickerby et al. teaches a traditional bond coat overlying a metallic substrate bonded to a TBC. The TBC comprises a plurality of layers, each layer having columnar grains, the columnar grains in each layer extending substantially perpendicular to the interface between the bond coating and metallic substrate. The structure is columnar to ensure that the strain tolerance of the ceramic TBC is not impaired. The difference in structure of the layers is the result of variations in the microstructure and/or density/coarseness of the columnar grains of the ceramic. These differences assist in providing to resistance to the transfer of heat across the thermal barrier coating.
Improved environmental resistance to destructive oxidation and hot corrosion is desirable. Additionally, the alloying elements of the bond coat interdiffuse with the substrate alloy at elevated temperatures of operation, changing the composition of the protective outer layer. Over time, as the airfoils are refurbished, walls of the airfoils are consumed, which reduces load carrying capability and limits blade life. Also, this interdiffusion can also reduce environmental resistance of the coating, causing loss of material, as layers of material are lost due to corrosive and oxidative effects. This interdiffusion and its adverse effects can be reduced by controlling the temperature of the component in the region of the bond coat/substrate interface. However, even with the use of advanced cooling designs and thermal barrier coatings, it is also desirable to decrease the requirement for cooling; because reducing the demand for cooling is also well known to improve overall engine operating efficiency.
One efficient cooling technique is film cooling. Film cooling is achieved by passing cooling air through discrete film cooling holes, typically ranging from 0.015xe2x80x3 to about 0.030xe2x80x3 in hole diameters. The film cooling holes are typically drilled with laser or by electro-discharge machining (EDM) or electro-stem (ES) machining. Due to mechanical limitations, each film hole has an angle ranging from 20xc2x0 to 90xc2x0 relative to the external surface. Therefore, each film jet exits from the hole with a velocity component perpendicular to the surface. But, because of this vertical velocity component and a complex aerodynamic flow circulation near the tip of a turbine blade, commonly referred to as the xe2x80x9csquealer tipxe2x80x9d, each film jet will have a tendency to lift or blow off from the external surface and mix with the hot exhaust gases, resulting in poor film cooling effectiveness in the area surrounding the squealer tip.
Thus, there is an ongoing need for an improved thermal barrier coating system surrounding the squealer tip, wherein the environmental resistance and long-term stability of the thermal barrier coating system is improved so that higher engine efficiencies can be achieved. The bond coat temperature limit is critical to the TBC""s life and has an upper limit of about 2100xc2x0 F. Once the bond coat exceeds this temperature, the coating system tends to quickly deteriorate, due to high temperature mechanical deformation and oxidation, as well as from interdiffusion of elements with the substrate alloy and subsequent degradation due to loss of its superior environmental resistance. The coating system can separate from the substrate exposing the underlying superalloy component to damage from the hot gasses.
In particular, the squealer tip is the most difficult location to cool in a turbine blade. The squealer tip is located away from the convection cooling in the center of the blade, and the complex aerodynamic flow field near the squealer tip makes film cooling very inefficient. This inefficient cooling results in tip deterioration much earlier than desired, and requires tip repairs after relatively short time in service to recover the tip clearance for better turbine efficiency.
As described above, to be effective, a TBC system requires active cooling on the backside of the location to be cooled. In the past, a TBC system was not used in the squealer tip region because, due to physical limitations, no effective backside cooling was available to make the TBC system effective. During the blade manufacturing process, to prevent the TBC from adhering to this area, the squealer tip and tip cap are usually masked during the TBC coating process. However, recently, in order to reduce manufacturing costs, the TBC coating application has been extended to cover the squealer tip region, thus avoiding the time-consuming masking process. It is therefore advantageous to further improve squealer tip cooling efficiency by intentionally incorporating an effective TBC coating system that includes backside cooling to extend squealer tip life.
In a co-pending application identified as Attorney Docket No. 13DV-13654 entitled xe2x80x9cCooled Thermal Barrier Coatings On a Turbine Blade Tipxe2x80x9d, channeled apertures were provided to a substrate in a preselected diameter sufficient to allow passage of cooling fluid from a cooling fluid source. A sufficiently thick bond coat was then applied to the substrate such that the bond coat partially filled the apertures. A porous TBC layer was then applied to the bond coat to partially occupy the apertures, followed by application atop the porous TBC layer of an additional dense ceramic TBC layer, denser than the porous TBC layer overlying the bond coat. Optionally, a conventional TBC system was applied on the concave (pressure side) and convex (suction side) of the airfoil surface.
In this manner, cooling fluid passes from a cooling fluid source through a channel aperture substantially adjacent to the squealer tip into the porous TBC. Because the channel aperture is at least partially filled with porous TBC material, cooling fluid flowed through the partially filled passageways. But, due to the increased resistance to flow, the cooling fluid expanded its flow path into the adjacent, porous TBC layer, continuing to flow between the bond coat and the dense coat, thereby directing cooling fluid to the squealer tip as the fluid exits the blade. However, because the porosity in the porous TBC layer is randomly distributed, there is little control of cooling fluid through the porous TBC and cooling can be irregular. What is needed are further improved designs that will allow a turbine engine blade squealer tip to run at higher operating temperatures, thus improving engine performance without the need for additional cooling air. This can be achieved by providing a regular cooling pattern through the porous TBC to the squealer tip. It is also desirable to have a system that can take advantage of the thermal insulation provided by TBC. The present invention fulfills this need, and further provides related advantages.
The present invention provides for cooling the squealer tip region of a high pressure turbine blade used in a gas turbine engine. The squealer tip is coated with a metallic bond coat and micro grooves are fabricated in an airfoil oriented in a substantially radial direction on the interior surface of the squealer tip above the tip cap. A micro groove oriented in a substantially axial direction is fabricated along or near the joint corner between the squealer tip side wall and the tip cap to provide fluid communication with all of the radial micro grooves. Tip cap cooling holes are drilled through the tip cap so as to be in fluid communication with the axial micro groove. TBC ceramic is then deposited by a shadowing technique on both blade external surfaces and the tip cavity, forming micro channels from the micro grooves.
In this manner, cooling fluid passes from a cooling fluid source through the tip cap holes and into a plenum created by the axially-oriented micro channel. The cooling fluid then passes from the plenum into the radially-oriented micro channels. Cooling fluid is thereby directed through the micro channels to cool the squealer tip by exiting to the tip. Since the TBC is porous, some of the cooling fluid will also provide transpiration cooling as it flows through the TBC.
The present invention further comprises both a cooled blade and a squealer tip region formed by the foregoing methods and techniques as well as the blade and squealer tip having micro channels for cooling the squealer tip.
One advantage of the present invention is that convection cooling through the micro channels inside or adjacent to a bond coat layer provides direct and efficient cooling for squealer tips.
Another advantage of this invention is that the TBC on the external surfaces of the squealer tip also becomes a very effective insulation. This insulating TBC has increased service life due to the backside cooling provided by the present invention. The combination of the effective convection cooling inside the micro channels and effective thermal insulation on the external surfaces of the squealer tip will lower the temperature of the squealer tip as compared to conventional and current designs, providing a longer service life. The beneficial cooling effects of the present invention create increased adherence of the pressure and suction side TBC to the airfoil component.
By removing heat from this region, the integrity of the bond coat also can be maintained at higher engine firing temperatures by inhibiting temperature-induced diffusion, resulting in a more efficient usage of cooling fluid than that of the prior art to achieve a higher turbine engine efficiency and performance while improving squealer tip service life.
Other features and advantages of the present invention will be apparent from the following more detailed description of the preferred embodiment, taken in conjunction with the accompanying figures which illustrate, by way of example, the principles of the invention.